1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to turbine airfoils with a cooling circuit.
2. Description of the Related Art Including Information Disclosed under 37 CFR 1.97 and 1.98
Gas turbine engines, especially of the axial flow type, include turbines to convert a hot gas stream from the combustor into mechanical energy. The turbine is formed of an alternating series of stationary vanes or nozzles followed by a stage or rotating blades or buckets. The first stage vane and blade arrangement is exposed to the highest gas stream temperature. The higher the temperature into the turbine, the higher the efficiency of the gas turbine engine. As the hot gas flow passes through the turbine, the gas flow temperature progressively decreases as the turbine converts the high temperature gas into mechanical energy. In order to allow for high gas flow temperature into the turbine, internal air cooling of the vanes and blades, especially in the first and second stages of the turbine, is required.
In the Prior Art, both blades and vanes are cooled internally with cooling air that has been bled off from the compressor of the gas turbine engine. Since the turbine drives the compressor, air that is bled off from the compressor and not burned with a fuel in the combustor lowers the overall efficiency of the engine. Thus, the larger the amount of air bled off from the compressor the lower the engine overall efficiency. Another design problem with internal cooled turbine airfoils, especially for the blades, is the effect of stress when combined with high temperature loads on the blades. The blades rotate at high speeds and therefore induce high centrifugal force that produces high stress levels especially at the root of the blade. The stress levels on the blade are lower near the tip region. For this reason, blades are design to have lower operating temperatures near the root that at the tip. Thus, more cooling is required near the root of the blade than near the tip on the airfoil surface.
The prior art is filled with various inventions for providing internal cooling air passages within airfoils used in the gas turbine engine. Serpentine cooling passages are an effective arrangement for providing cooling to the blade as well as limiting the amount of cooling air bled off from the compressor. U.S. Pat. No. 6,264,428 B1 issued to Dailey et al on Jul. 24, 2001 entitled COOLED AEROFOIL FOR A GAS TURBINE ENGINE discloses a hollow cooled turbine blade with internal cooling passages. The blade root includes a cooling air inlet passage (22 in this patent) leading into a plurality of suction side radial cooling air passages (21 in this patent), reverses flow into a central plenum (16 in this patent), and then passes through a plurality of apertures (27 in this patent) into a plurality of radial cooling air passages (28 in this patent) on the pressure side of the airfoil. The Dailey et al patent passes the cooling air into the radial channel on the suction side, which is exposed to a lower gas flow temperature than is the pressure side. Also, use of the turbulators in the channels increases the pressure loss as the cooling air passes through, requiring a higher pressure head on the cooling air flow through the airfoil.
Another prior art patent, U.S. Pat. No. 5,702,232 issued to Moore on Dec. 30, 1997 entitled COOLED AIRFOILS FOR A GAS TURBINE ENGINE discloses a turbine blade with a feed chamber (58 in this patent) connected to re-supply holes (59 in this patent) that lead into feed passages (57 and 56 in this patent), which lead into film holes (65 in this patent) that discharge cooling air onto the airfoil surface. Because the film cooling holes (65 in this patent) lead from the radial channels (56 and 57 of this patent), the openings of the film cooling holes are without diffusers, or the wall thickness must be increased to allow for the use of diffusers on the hole openings.
In order to improve on the efficiency of a gas turbine engine, the operating temperature can be increased which requires improved cooling of the airfoil. Also, the efficiency of the engine can be improved by using less bleed air from the compressor. It is therefore an object of the present invention to provide improved cooling for an airfoil of a gas turbine engine that uses internal cooling passages supplies by a flow of cooling air. It is another object of the present invention to provide for a cooling circuit that also requires less cooling flow to provide cooling for the airfoil. Another object of the present invention is to provide for a turbine blade with more cooling at the root of the blade than at the tip of the blade without increasing the amount of cooling air needed to cool the blade.